This invention relates generally to an assembly of turbine engine articles distinct from one another and disposed about rotating articles. One example includes a turbine shroud disposed about rotating blading members, in a series of associated, juxtaposed distinct members that can comprise a combination of a stationary nozzle with vanes, engine frames, etc.
It is typical in the turbine engine art, for example art relating to gas turbine engines, to dispose a series of generally stationary members radially outwardly from an engine axis of rotation about rotating blades to define together a part of a radially outer flowpath boundary of the engine. An example of such a series of members, axially extending in the engine and juxtaposed one with another, comprises a turbine stator or nozzle having a stage of vanes each including an outer band; a turbine shroud circumferentially about rotating turbine blades; and a turbine engine rear frame or another turbine nozzle. In many assemblies, axially adjacent members of such a series are in juxtaposition across an axial gap that requires a separate fluid seal to inhibit the radially outward flow of the engine gas stream and/or the radially inward flow of cooling air. As is well known in the gas turbine engine art, engine efficiency can be reduced by fluid losses resulting from leakage through such gaps. Some examples of U.S. Patents relating to such structures include U.S. Pat. No. 5,071,313xe2x80x94Nichols; U.S. Pat. No. 5,074,748xe2x80x94Hagel; U.S. Pat. No. 5,127,793xe2x80x94Walker et al.; and U.S. Pat. No. 5,562,408xe2x80x94Proctor et al.
Metallic type materials currently and typically are used to make members in such a series, including shrouds and shroud segments. Therefore, some engine assemblies include a series of metallic members, such as a series of stationary nozzle vanes, shrouds, and/or frames and other vanes, in contact with each other and axially loaded together to define a substantially continuous flowpath portion in the engine. One such example is shown in U.S. Pat. No. 3,807,891xe2x80x94McDow et al. That kind of loading or restraint can result in the application of a substantial compressive force to the members. If such members are made of typical high temperature alloys generally currently used in gas turbine engines, the alloys can easily withstand and accommodate such compressive forces. However, if one or more of the series of members is made of a low ductility, relatively brittle material, such compressive loading can result in fracture or other detrimental damage to the member during engine operation.
Current gas turbine engine development has suggested, for use in higher temperature applications such as shroud segments and other components, certain materials having a higher temperature capability than the metallic type materials currently in use. However such materials, forms of which are referred to commercially as a ceramic matrix composite (CMC), have mechanical properties that must be considered during design and application of an article such as a shroud segment. For example, as discussed below, CMC type materials have relatively low tensile ductility or low strain to failure when compared with metallic materials. Also, CMC type materials have a coefficient of thermal expansion (CTE) in the range of about 1.5-5 microinch/inch/xc2x0 F., significantly different from commercial metal alloys used as restraining supports or hangers for shrouds of CMC type materials. Such metal alloys typically have a CTE in the range of about 7-10 microinch/inch/xc2x0 F. Therefore, if a CMC type of shroud segment is restrained or axially loaded with an offset reaction point during engine operation, and cooled on one surface as is typical during operation, compressive forces can be developed in a CMC type segment sufficient to cause failure of the segment.
Generally, commercially available CMC materials include a ceramic type fiber for example SiC, forms of which are coated with a compliant material such as BN. The fibers are carried in a ceramic type matrix, one form of which is SiC. Typically, CMC type materials have a room temperature tensile ductility of no greater than about 1%, herein used to define and mean a low ductility material. Generally CMC type materials have a room temperature tensile ductility in the range of about 0.4-0.7%. This is compared with metallic shroud and/or supporting structure or hanger materials having a room temperature tensile ductility of at least about 5%, for example in the range of about 5-15%. Shroud segments made from CMC type materials, although having certain higher temperature capabilities than those of a metallic type material, cannot tolerate the above described and currently used type of compressive force or similar restraint force against chording. Therefore, a shroud segment assembly, in one embodiment including shroud segments of a low ductility material, floating axially independently of other engine members and positioned or disposed in a manner that does not apply detrimental force to the shroud segment during operation enables advantageous use of the higher temperature capability of CMC material. Provision of a turbine engine series of members including an intermediate member axially floating independently of adjacent members and separated prior to engine operation from an axially aft member across a selected gap can enable axial sealing of the assembly without additional seal members and without application of excessive loading or a compressive force on the intermediate member by selective axially movement of the axially floating member. This can enable successful use of a CMC material for making a member such as a shroud or shroud segment and can eliminate or at least reduce the requirement for additional, separate seals.
One form of the present invention provides a combination of an axially disposed series of members in a turbine engine. The engine comprises a compressor section for compressing incoming fluid, a combustion section for burning fuel with the fluid to generate products of combustion or combustion gases, and a turbine section for extracting energy from the products of combustion. Each of the axially disposed series of members is axially distinct from an adjacent juxtaposed member at least at a radially outer portion. The combination comprises a series of three respectively juxtaposed members. One is an axially forward first member, for example a non-rotating nozzle. A second is an axially middle or intermediate second member, for example a shroud or shroud segment, floating independently axially. A third is an axially aft third member, for example another non-rotating nozzle or a portion of a turbine aft frame, separated prior to engine operation from the second member by a gap.
The first member includes a radially outer portion having an axially aft surface, and a radially inner portion held by the engine. Thus the radially outer portion is cantilevered from its radially inner portion, the axially aft surface of the first member being free to move a first axial length or movement distance axially aft as a result of typical aeronautical force or load applied to the first member during engine operation. Such axial movement of the first member radially outer portion reduces any first gap of first gap axial length that may exist after assembly between such portion and a juxtaposed portion of the second member. In addition, such movement applies force to move the axially floating second member a second axial length or movement distance. If substantially no gap exists between the first and second members, the second axial length is substantially the same as the first axial length.
The second member, floating independently axially of the other members, includes an axially forward surface in juxtaposition and for registry with the axially aft surface of the first member, and an axially aft surface. As assembled prior to engine operation, the aft surface of the first member and the forward surface of the second member can be in contact or can be separated by an axial gap, as mentioned above.
The third member includes an axially forward surface disposed prior to engine operation axially across a second gap, of pre-selected second gap axial length, with the second member axially aft surface. The second gap axial length of the second gap is selected, prior to engine operation, as a function of the first axial length or movement distance to substantially close any gaps between the second and third members during engine operation.
Another form of the present invention provides a method of assembling a turbine engine including the above-described series of members to provide at least the gap between the second and third members of the third axial length.